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  • 1
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    Effect of the NACA Injection Impeller on the Mixture Distribution of a Double-row Radial Aircraft Engine (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: For the normal range of engine power the impeller provided marked improvement over the standard spray-bar injection system. Mixture distribution at cruising was excellent, maximum cylinder temperatures were reduced about 30 degrees F, and general temperature distribution was improved. The uniform mixture distribution restored the normal response of cylinder temperature to mixture enrichment and it reduced the possibility of carburetor icing, while no serious loss in supercharger pressure rise resulted from injection of fuel near the impeller outlet. The injection impeller also furnished a convenient means of adding water to the charge mixture for internal cooling.
    Keywords: Aerodynamics
    Type: NACA-TN-1069
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Correlation Tests of the Ditching Behavior of an Army B-24D Airplane and a 1/16-size Model (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Behaviors of both model and full-scale airplanes were ascertained by making visual observations, by recording time histories of decelerations, and by taking motion picture records of ditchings. Results are presented in form of sequence photographs and time-history curves for attitudes, vertical and horizontal displacements, and longitudinal decelerations. Time-history curves for attitudes and horizontal and vertical displacements for model and full-scale tests were in agreement; maximum longitudinal decelerations for both ditchings did not occur at same part of run; full-scale maximum deceleration was 50 percent greater.
    Keywords: Aerodynamics
    Type: NACA-WR-L-617 , NACA-MR-L6A03
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    The Knock-Limited Performance of Fuel Blends Containing Spiropentane, Methylenecyclobutane, Di-Tert-Butyl Ether, Methyl Tert-Butyl Ether, and Triptane (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests show that at inlet-air temperatures of 250 deg F and 100 deg F the knock-limited performance of the base fuel of blends, leaded with 4 ml TEL per gallon and containing 20 percent spiropentane, was reduced at fuel/air ratios below 0.085. The 20 percent methylenecyclobutane reduced the knock-limited power of the base fuel at fuel/air ratios below 0.112. Di-tert-butyl ether, methyl-tert-butyl ether, and triptane increased the knock-limited power of the base fuel at all fuel/air ratios and at both temperatures.
    Keywords: Aerodynamics
    Type: NACA-WR-E-222 , NACA-RB-E6D22
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Flight Investigation to Improve the Dynamic Longitudinal Stability and Control-Feel Characteristics of the P-63A-1 Airplane (AAF No. 42-68889) with Closely Balanced Experimental Elevators (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results of flight tests of a control-feel aid presented. This device consisted of a spring and dashpot connected in series between the control stick and airplane structure. The device was tested in combination with an experimental elevator and bobweight which had given unsatisfactory dynamic stability and control-feel characteristics in previous tests. The control-feel aid effected marked improvement in both the control-feel characteristics and the control-feel dynamic longitudinal stability of the airplane.
    Keywords: Aerodynamics
    Type: NACA-WR-L-730 , NACA-MR-L6E20
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Field of Flow About a Jet and Effect of Jets on Stability of Jet-Propelled Airplanes (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A theoretical investigation was conducted on jet-induced flow deviation. Analysis is given of flow inclination induced outside cold and hot jets and jet deflection caused by angle of attack. Applications to computation of effects of jet on longitudinal stability and trim are explained. Effect of jet temperature on flow inclination was found small when thrust coefficient is used as criterion for similitude. The average jet-induced downwash over tail plane was obtained geometrically.
    Keywords: Aerodynamics
    Type: NACA-WR-L-213 , NACA-ACR-L6C13
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    High-Speed Wind-Tunnel Tests of a Twin-Fuselage Pursuit Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: At the request of the Air Technical Service Command, U.S. Army Air Forces, a 0.22-scale model of a twin-fuselae pursuit airplane was built and tested at the Ames Aeronautical Laboratory. The tests of this model were made in order that the aerodynamic characteristics of the airplane, especially at high speed, might be predicted. The results shown in this report consist of force data for the model and critical Mach numbers of parts of the model as determined from pressure-distribution measurements. The results indicate that a diving tendency of the airplane can be expected at Mach numbers above 0.70 at lift co-efficients from 0 to 0.4. There is an indication that the Mach number at which the airpolane would first experience a diving tendency for lift coefficients from 0 to 0.2 can be increased if the critical speed of the radiator enclosures is increased, and the wing-fuselage-juncture fillets are improved.
    Keywords: Aerodynamics
    Type: NACA-WR-A-75 , NACA-MR-A6D03
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.
    Keywords: Aerodynamics
    Type: NACA-WR-L-98 , NACA-ACR-L6B21
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Comparison of Wind-Tunnel and Flight Measurements of Stability and Control Characteristics of a Douglas A-26 Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests in Langley pressure tunnel of model XA-26 bomber were compared with those of A-26B (twin-engine attack bomber) and showed that static longitudinal stability, indicated by elevator-fixed neutral points, and variation of elevator deflection in straight and turning flight were good. Airplane possessed improved stability at low speeds which was attributed to pronounced stalling at root of production wing. At rudder-force reversal at speeds higher than those in flight tests, agreement in rudder-fixed and rudder-free static directional stability was good. Hinge moment obtained at zero sideslip was satisfactory for determining aileron forces in sideslip.
    Keywords: Aerodynamics
    Type: NACA-WR-L-99 , NACA-ARR-L5H11a
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Analysis of V-G Records from the SNB-1 Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Availability data obtained on SNB-1 trainer-class airplanes were analyzed and results presented as flight envelopes which predict occurrences of large values of air speed and acceleration. Comparison is made with SNJ-4 trainer-class airplane data analyzed by the same method. It is concluded that flight envelopes are satisfactory; that the two types show large differences in flight loads and speeds experience; and that SNB-1 will seldom, if ever, exceed design limit load factor and restricted speed, which SNJ-4 can be expected to exceed design-limit load factor and restricted speed in a very small number of flight hours.
    Keywords: Aerodynamics
    Type: NACA-WR-L-759 , NACA-MR-L6F27a
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    The Knock-Limited Performance of Fuel Blends Containing Aromatics V : N-Propylbenzene, N-Butylbenzene, Isobutylbenzene, M-Xylene, and 1-Isopropyl-4-Methylbenzene (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results are reported of knock-limited tests of five aromatics, each individually blended with selected base fuels and tested with and without TEL, using 17.6, F-4, and F-3 small-scale engines. The five aromatics rated in the following order of decreasing antiknock effectiveness at fuel/air ratio 0.10: m-xylene, 1-isopropyl-4-methylbenzene, n-propylbenzene, isobutylbenzene, and n-butylbenzene.
    Keywords: Aerodynamics
    Type: NACA-WR-E-237 , NACA-ARR-E6C05
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 11
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    An Experimental Investigation of Flow Conditions in the Vicinity of an NACA D(sub S)-type Cowling (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Data are presented of the flow conditions in the vicinity of an NACA D sub S -type cowling. Tests were made of a 1/2 scale-nacelle model at inlet-velocity ratios ranging from 0.23 to 1.02 and angles of attack from 6 deg to 10 deg. The velocity and direction of flow in the vertical plane of symmetry of the cowling were determined from orifices and tufts installed on a board aligned with the flow. Diagrams showing velocity ratio contours and lines of constant flow angles are given.
    Keywords: Aerodynamics
    Type: NACA-WR-L-747 , NACA-MR-L6H14
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 12
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    Development of Wing Inlets (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil.
    Keywords: Aerodynamics
    Type: NACA-WR-L-727 , NACA-ACR-L6B18
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 13
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    Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.
    Keywords: Aerodynamics
    Type: NACA-WR-W-2 , NACA-ACR-6A30
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 14
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    Analysis of Effect of Rolling Pull-Outs on Wing and Aileron Loads of a Fighter Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analysis was made to determine the effect of rolling pull-out maneuvers on the wing and aileron loads of a typical fighter airplane, the P-47B. The results obtained indicate that higher loads are imposed upon wings and ailerons because of the rolling pull-out maneuver, than would be obtained by application of the loading requirements to which the airplane was designed. An increase of 102 lb or 15 percent of wing weight would be required if the wing were designed for rolling pull-out maneuver. It was also determined that the requirements by which the aileron was originally designed were inadequate.
    Keywords: Aerodynamics
    Type: NACA-WR-L-270
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 15
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    Free-Spinning Tunnel Tests of a 1/20-Scale Model of the Chance Vought XF6U-1 Airplane, TED No. NACA 2390 (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A spin investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/20-scale model of the Chance Vought XF6U-1 airplane, The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the normal-fighter condition. The investigation also included tests for the take-off fighter condition (wing-tip tanks plus fuel added) spin-recovery parachutes, and simulated pilot escape. In general, for the normal-fighter condition, the model was extremely oscillatory in roll, pitch, and yaw. The angles of the fuselage varied from extremely flat to inverted attitudes, and the model rotated with the rudder in a series of short turns and glides. Recoveries by rudder reversal were rapid but the model would immediately go into a spin in the other direction. Recoveries by merely neutralizing the rudder were satisfactory when the elevator and ailerons were set to neutral, the ensuing flight path being a steep glide. Thus, it is recommended that all controls be neutralized for safe recovery from spins obtained on the airplane. With the external wing-tip tanks installed, the spins were somewhat less oscillatory in roll but recovery could not be obtained unless full-down elevator was used in conjunction with the rudder. If a spin is entered inadvertently with the full-scale airplane with external wing-tip tanks installed and if recovery is not imminent after a recovery attempt is made, it is recommended that the tanks be jettisoned and the controls neutralized.
    Keywords: Aerodynamics
    Type: NACA-RM-L6H27
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 16
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    Free-Spinning-Tunnel Tests of a 1/20-Scale Model of the Northrop N-9M Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Spin tests of a 1/20-scale model of the Northrop N-9M airplane have been performed in the Langley 20-foot free-spinning tunnel. The erect and inverted spin and recovery characteristics were determined for various loading conditions and the effect of deflecting the flaps and of extending the landing gear was investigated. The investigation also included tests to determine the size parachute required for satisfactory spin recovery by parachute action alone. The tests were performed at an equivalent spin altitude of 15,000 feet. A specialized recovery technique consisting of rapid full reversal of the rudder pedals against the spin combined with turning the wheel against the spin and movement of the stick forward is recommended for all loadings and configurations of the airplane. The results also indicated that a 7-foot-diameter spin-recovery parachute having a drag coefficient of 0.7 attached to the outboard wing tip with a towline of 10 to 30 feet or an 8.8-foot-diameter parachute attached to the fixed portion of the wing between the elevons and the pitch flaps with a 30-foot towline would provide satisfactory recovery from demonstration spins by parachute action alone. It appears possible that the first N-9M airplane may have crashed because of failure to recover from a spin.
    Keywords: Aerodynamics
    Type: NACA-RM-L6G30
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 17
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    Calibration Tunnel for High Speed (1946)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: For the nvestigation of measuring instruments at higher speeds up to a Mach number 0.7 a tunnel with closed test section was built in 1942 which was as simple and cheap as possble. The blower was a radial blower with straight sheet vanes of 800-millimeter diameter the tips of which were bent backward a little. The blower sucks the air through a honeycomb of diameter 1.2 neter with wide meshes. The air is then accelerated in a short cone with smooth transition to the test section. The cylindrical test section of 200-milimeter diameter has two windows (which are displaced 180 deg from each other. The instruments may be introduced and observed through and observed through these windows. . The cross section is then enlarged by a straight diffuser 3.5 meters long and reaches the ninefold cross section. The air flows back into the room through a disk diffuser of 2-meter diameter. The maximum speed in the jet is 250 m/s for a drive power of 35 kT., if there are no installations in the jet. The velocity is determined by pressure holed along the test section.
    Keywords: Aerodynamics
    Type: NACA/TM-1103
    Format: application/pdf
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  • 18
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    Theoretical and Experimental Data for a Number of NACA 6A-Series Airfoil Sections (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The NACA 6A-series airfoil sections were designed to eliminate the trailing-edge cusp which is characteristic of the NACA 6-series sections. Theoretical data are presented for NACA 6A-series basic thickness forms having the position of minimum pressure at 30-, 40-, and 50-percent chord and with thickness ratios varying from 6 percent to 15 percent. Also presented are data for a mean line designed to maintain straight sides on the cambered sections. The experimental results of a two dimensional wind tunnel investigation of the aerodynamic characteristics of five NACA 64A-series airfoil sections and two NACA 63A-series airfoil sections are presented. An analysis of these results, which were obtained at Reynolds numbers of 3 x 10(exp 6), 6 x 10(exp 6), and 9 x 10(exp 6), indicates that the section minimum drag and maximum lift characteristics of comparable NACA 6-series and 6A-series airfoil sections are essentially the same. The quarter-chord pitching-moment coefficients and angles of zero lift of NACA 6A-series airfoil sections are slightly more negative than those of corresponding NACA 6-series airfoil sections. The position of the aerodynamic center and the lift-curve slope of smooth NACA 6-series sections. The addition of standard leading-edge roughness causes the lift-curve slope of the newer sections to decrease with increasing airfoil thickness ratio.
    Keywords: Aerodynamics
    Type: NACA-RM-L6J01
    Format: application/pdf
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  • 19
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    Plane and Three-Dimensional Flow at High Subsonic Speeds (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: For two- and three-dimensional flow in a compressible medium, a simple relation is given by which, to a first approximation, the quantitative influence of compressibility upon the velocities and pressures can be understood in a clear manner. In the application of this relation the distinct behaviors of two-dimensional and axially symmetric three-dimensional flow with increasing Mach number are brought out. For slender elliptic cylinders and ellipsoids of revolution, calculations are made of the critical Mach number; that is, the Mach number at which local sonic velocity is achieved on the body. As a further example, the lifting wing of finite span is considered, and it is shown that the increase of wing lift with Mach number at a given angle of attack is greatly dependent upon the aspect ratio b(exp 2)/F.
    Keywords: Aerodynamics
    Type: NACA-TM-1105 , Lilienthal Gesellschaft; 127; 97-101
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 20
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    Wind-Tunnel Tests of a Modified Koppers Aeromatic Impeller-Generator Combination, TED No. NACA ARR 2901 (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation was conducted in the 6- by 6-foot test section of the Langley stability tunnel on a modified Koppers Aeromatic wind-driven impeller-generator combination. This investigation consisted of a few fixed pitch tests and a series of variable pitch tests, The fixed pitch tests indicated that the impeller should operate between the blade-pitch angles of 20 and 32deg at the specified output of 11.7 kilowatts in order to remain within the specified rotational speed of from 5000 to 8000 rpm for airspeeds of from 130 to 175 miles per hour. The requirement that the impeller maintain rotational speeds of - between 5000 and 8000 rpm as the impeller output varied from 0 to ll.7 kilowatts at airspeeds of from 130 to 175 miles per hour was not met at any tims during the variable pitch tests. The main difficulty seemed to be the inability of the impeller blades to change blade-pitch angle smoothly and quickly as load conditions varied. There was some indication that the vibration normally occurring on an airplane might cause the impeller to operate satisfactorily.
    Keywords: Aerodynamics
    Type: NACA-RM-L6K19
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 21
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    Wind-Tunnel Investigation of the Stability of the Jettisonable Nose Section of the X-3 Airplane (1946)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Because previous work has indicated that jettisonable nose sections of airplanes may be inherently unstable, and thus may cause dangerous centripetal accelerations on a pilot after jettisoning during high-speed flight, an investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the behavior in descent of a model of the jettisonable nose section of the Douglas X-3 airplane. The effects of varying the center-of-gravity position, of attaching fins of various sizes, and of installing a stabilizing parachute were investigated. In the investigation the model descended with its front and trimmed 36 deg above the horizontal and rotated about a vertical wind axis while rolling about its longitudinal body axis. The nose section was made to descend in a stable front-down attitude when stabilizing fins were installed in conjunction with movement of the center of gravity forward or when a stable parachute was attached to the model.
    Keywords: Aerodynamics
    Type: NACA-RM-L8K18a
    Format: application/pdf
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